This invention relates to composite materials and, more particularly, to repairing composite sandwich structures.
Composite structures are commonly known and used in, for example, gas turbine engines for aerodynamic fairings. One type of composite is a sandwich structure having a face skin and a back skin that are bonded to a honeycomb core. The face skin and back skin may be made of a metal sheet or a laminated structure with a resin matrix, such as epoxy, with one or more plies of reinforcement fibers. Often, the face skin, the back skin, or both are perforated for noise attenuation.
In some instances, it is desirable to repair a section of the face skin that has become damaged from an impact, elevated temperatures, abrasion, erosion, or other phenomenon rather than replace the entire face skin. A typical repair for this includes using a potting material to restore the aerodynamic profile of the face skin and provide load transfer across the damaged area. However, the potting material blocks some of the perforations, thereby decreasing the acoustic performance.
Alternatively, if the face skin is a laminated structure, one or more laminate plies may be removed and patched using what is known as a “scarf.” Often, an oversized patch is placed over the repair area and is sized to create a lap joint in the area adjacent to the repair. The “scarf” repair method has a significant drawback in that the patch may peel off of the repair area or entirely liberate under extreme conditions.
Additionally, one trend has been to make the face skin as thin as possible to reduce weight and enhance acoustics. The thinness makes it difficult to use the traditional “scarf” repair method because there may only be one or two laminate plies. Furthermore, the “scarf” repair method is not even applicable for metal face skins. Thus, there is a need for a reliable repair method for laminated composite and metal face skins. This invention addresses those needs while avoiding the shortcomings and drawbacks of the prior art.